A gas turbine engine may be used to supply power to various types of vehicles and systems. For example, gas turbine engines may be used to supply propulsion power to an aircraft. Many gas turbine engines include at least three major sections, a compressor section, a combustor section, and a turbine section. The compressor section receives a flow of intake air and raises the pressure of this air to a relatively high level. In a multi-spool (e.g., multi-shaft) engine, the compressor section may include two or more compressors. The compressed air from the compressor section then enters the combustor section, where a ring of fuel nozzles injects a steady stream of fuel. The injected fuel is ignited by a burner, which significantly increases the energy of the compressed air.
The high-energy compressed air from the combustor section then flows into and through the turbine section, causing rotationally mounted turbine blades to rotate and generate energy. The air exiting the turbine section is then exhausted from the engine. Similar to the compressor section, in a multi-spool engine the turbine section may include a plurality of turbines. The energy generated in each of the turbines may be used to power other portions of the engine.
In addition to providing propulsion power, a gas turbine engine may also, or instead, be used to supply either, or both, electrical and pneumatic power to the aircraft. For example, some gas turbine engines include a bleed air port between the compressor section and the turbine section. The bleed air port allows some of the compressed air from the compressor section to be diverted away from the turbine section, and used for other functions such as, for example, the aircraft environmental control system, and/or cabin pressure control system.
Although highly unlikely, it is postulated that a leak may occur in one or more of the ducts through which engine bleed air is flowing. In such an unlikely circumstance, bleed air flow rate through the leaking duct can be detrimental to turbine engine operation. Thus, many bleed air ducts include orifices or bleed dumps, which limit flow through the duct, if a leak were to occur. While these present flow limiting configurations do limit flow, these configurations also suffer certain drawbacks. For example, the present flow limiting configurations can result in undesirably large pressure losses to achieve a desired bleed air flow during normal bleed air operations, which can adversely impact gas turbine engine efficiency. Moreover, bleed air may need to be supplied from higher pressure stages of the compressor section, or designing compressor stages to a higher pressure ratio. Bleeding air from higher pressure stages can increase fuel burn and cause higher bleed air temperatures. This in turn can adversely impact overall operational efficiency and costs.
Hence, there is a need for a system for limiting flow in bleed air ducts that, as compared to present systems, exhibits reduced pressure losses during normal bleed air operations and/or does not adversely impact gas turbine engine efficiency and/or does not adversely impact overall operational efficiency and cost. The present invention addresses one or more of these needs.